|Publication number||US6790005 B2|
|Application number||US 10/334,091|
|Publication date||Sep 14, 2004|
|Filing date||Dec 30, 2002|
|Priority date||Dec 30, 2002|
|Also published as||US20040126236|
|Publication number||10334091, 334091, US 6790005 B2, US 6790005B2, US-B2-6790005, US6790005 B2, US6790005B2|
|Inventors||Ching-Pang Lee, David Glenn Cherry, Chander Prakash, Aspi Rustom Wadia|
|Original Assignee||General Electric Company|
|Export Citation||BiBTeX, EndNote, RefMan|
|Patent Citations (16), Non-Patent Citations (1), Referenced by (90), Classifications (11), Legal Events (6)|
|External Links: USPTO, USPTO Assignment, Espacenet|
The present invention relates generally to gas turbine engines, and, more specifically, to turbine blade cooling therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases which flow downstream through turbine stages that extract energy therefrom. A high pressure turbine powers the compressor, and a low pressure turbine powers an upstream fan in a turbofan aircraft engine embodiment.
The first stage turbine blades first receive the hot combustion gases from the combustor and are typically air cooled by using air bled from the compressor. Turbine blade cooling is quite esoteric and the art is well crowded in view of the complex nature of blade cooling.
A typical turbine blade includes a generally concave pressure side and an opposite, generally convex suction side extending axially or chordally between leading and trailing edges which extend radially in span from root to tip of the blade. The airfoil portion of the blade is hollow and extends radially outwardly from a supporting dovetail which mounts the blade in a supporting rotor disk.
Cooling air is channeled to each blade through the dovetail and various internal passages are formed inside the airfoil for tailoring cooling thereof to mitigate the various heat loads experienced around the outer surface of the airfoil.
The radially outer end or tip of the airfoil is particularly difficult to cool since it is exposed to hot combustion gases along both the pressure and suction sides of the airfoil as well as in the radial clearance or gap formed with the surrounding stator casing or shroud. Since turbine blades are subject to occasional tip rubs, the airfoil tip is typically formed by squealer rib extensions of the pressure and suction sides which join together at the leading and trailing edges and define an open tip plenum therebetween having a floor which encloses the internal passages of the airfoil.
A significant advancement in blade tip cooling is U.S. Pat. No. 5,261,789 which discloses the use of a tip shelf along the pressure side of the turbine blade. The tip shelf is fed with cooling air through holes formed therethrough and interrupts the flow of combustion gases along the pressure side of the blade tip. Improved cooling of the blade tip including the pressure side tip rib is obtained.
During operation, combustion gases flow axially over the pressure and suction sides of the airfoil, with a portion thereof migrating radially upwardly along the pressure side and over the pressure side tip rib where it leaks past the airfoil tip in the small gap formed with the shroud. The resulting flow field of the combustion gases and cooling air discharged from the tip shelf is complex and affects both aerodynamic performance of the airfoil and cooling of the tip ribs themselves which are solid members extending upwardly from the tip floor.
Although the tip shelf and ribs are relatively small features of the airfoil, the importance thereof cannot be overstated since oxidation of the tip and material lost therefrom limits the useful life of the blade. The tip ribs are typically manufactured by casting with the entirety of the blade itself, and the small tip shelf may also be formed by casting or by electrical discharge machining (EDM) where required or practical. In either manufacturing method, the pressure side tip rib and cooperating tip shelf have dimensions measured in several mils, and are thus subject to manufacturing tolerances which affect the performance thereof.
Furthermore, the individual tip ribs are subject to centrifugal loading during operation which generates corresponding stress at the bases thereof with the tip floor. And, the tip shelf joins the pressure side tip rib at a correspondingly small fillet at which centrifugal stress may be concentrated during rotation of the blades in operation.
Accordingly, it is desired to provide a turbine blade having improved tip cooling notwithstanding manufacturing tolerances and centrifugal loads.
A gas turbine engine blade includes pressure and suction sidewalls extending between leading and trailing edges and from root to tip. The pressure sidewall includes an inclined tip rib offset therein by a ramp defining a tip notch having compound inclinations.
The invention, in accordance with preferred and exemplary embodiments, together with further objects and advantages thereof, is more particularly described in the following detailed description taken in conjunction with the accompanying drawings in which:
FIG. 1 is an isometric view of a gas turbine engine first stage rotor blade having blade tip cooling in accordance with an exemplary embodiment of the present invention.
FIG. 2 is a radial sectional view through a portion of the blade airfoil illustrated in FIG. 1 and taken generally along line 2—2.
FIG. 3 is an elevational sectional view through the airfoil illustrated in FIG. 1 and taken along line 3—3.
Illustrated in FIG. 1 is an exemplary first stage turbine rotor blade 10 for a gas turbine engine over which is channeled hot combustion gases 12 during operation. The blade includes a hollow airfoil 14 integrally joined to a mounting dovetail 16 typically formed in a common casting. The airfoil is configured for extracting energy from the combustion gases. And, the dovetail is configured for securing the blade in the perimeter of a rotor disk (not shown) which is rotated during operation.
As shown in FIGS. 1 and 2, the airfoil has a crescent-shaped aerodynamic profile including opposite pressure and-suction sidewalls 18,20 which extend axially or chordally between opposite leading and trailing edges 22,24. The airfoil sides also extend radially in longitudinal span from a root 26 at an integral blade platform 28 to a radially outer tip 30.
The hollow airfoil may have any conventional internal cooling circuit 32 and typically includes multiple internal flow passages having suitable inlets in the dovetail 16 through which cooling air 34 is received from a compressor (not shown) of the engine for use as a coolant in cooling the blade during operation.
As shown in FIG. 2, there are nine internal flow passages extending radially inside the airfoil for preferentially cooling the various portion thereof. The first two passages at the airfoil leading edge provide dedicated impingement cooling of the leading edge. The two passages in front of the trailing edge 24 provide dedicated cooling of the thin trailing edge region of the airfoil. And the five intermediate passages are arranged in a five-pass serpentine circuit for cooling the middle of the airfoil.
The airfoil includes various radial rows of film cooling holes 36 through the pressure and suction sidewalls of the airfoil as required for providing effective cooling thereof. And, the airfoil includes a row of trailing edge discharge holes 38 for discharging the cooling air from the last two internal passages axially outwardly along the trailing edge.
As shown in FIG. 3, a generally flat tip floor 40 bridges the airfoil pressure and suction sidewalls and encloses the several internal passages therein. The tip floor is bounded by integral first and second squealer tip ribs 42,44 extending along the pressure and suction sidewalls, respectively, and joining each other at the leading and trailing edges of the airfoil. The pressure and suction side ribs 42,44 surround the blade tip and extend upwardly from the floor to define an open tip cavity or plenum 46.
As best illustrated in FIG. 3, the pressure side or first squealer rib 42 is inclined outwardly from the tip plenum at an acute inclination angle A relative to the tip floor or horizontal axis, and has a base offset or recessed inwardly from the pressure sidewall 18 to form an exposed outboard inclined shelf or ramp 48. The ramp extends from the pressure sidewall to the first rib and defines a substantially normal notch 50 therebetween. Since both the first rib and the ramp are inclined to define the notch 50, the notch has compound angles of inclination, with the two legs thereof having a generally L-shaped configuration and are preferably orthogonal or normal to each other.
Since the first rib 42 is inclined from the tip floor, it bounds the pressure side of the tip plenum 46 at an obtuse included angle B between the corresponding surfaces thereof.
Correspondingly, the tip ramp 48 is inclined upwardly at an acute inclination angle C from an arcuate bullnose 52 at the juncture with the pressure sidewall 18 to join the first rib at an arcuate fillet 54. The bullnose provides a smooth external corner between the ramp and pressure sidewall, and the fillet provides a smooth internal corner with the first rib. And, the compound notch 50 extends chordally along the first rib 42 over a suitable distance between the leading and trailing edges of the airfoil as illustrated in FIG. 1. The notch decreases in size and blends as it terminates at suitable locations near the leading and trailing edges.
The airfoil tip illustrated in FIGS. 1 and 3 also includes a row of film cooling apertures 56 extending from the internal cooling circuit 32 through the juncture of the pressure sidewall and tip floor to reach the ramp 48 for providing cooling air thereto during operation.
The compound angled, pressure side, tip notch 50 may be used to provide substantial advantages in the aerodynamic and structural performance of the turbine blade, and in the casting manufacture thereof. The cooling air is discharged from the internal cooling circuit through the ramp apertures 56 to fill with film cooling air the tip notch 50 along the outboard surface of the first rib 42. The air in the notch can create a strong recirculation zone and provide enhanced thermal protection of the first rib 42.
Furthermore, the inclination of the first rib 42 causes the spent cooling air from the notch to be discharged over the top of the first rib as a barrier to the combustion gas flow thereover which in turn is bounded by the radially outer shroud (not shown). The obtuse turning angle of the cooling air around the pressure side rib 42 provides another recirculation zone over the top of the first rib which aerodynamically restrains passage of the combustion gases through the blade tip gap. In this way, efficiency of the turbine may be improved.
In the preferred embodiment illustrated in FIG. 3, the ramp 48 is substantially straight in cross section from the bullnose 52 to the fillet 54 and is disposed higher in elevation than the tip floor 40 inside the tip plenum. The two squealer ribs 42,44 have a preferred height measured from the top of the tip floor 40 conventionally determined to minimize the possibility of blade tip rubbing during operation while ensuring the structural integrity of the squealer ribs for minimizing combustion gas leakage through the tip gap. Since the ramp 48 is higher than the tip floor it reduces the radial distance from the outlet of the ramp apertures 56 to the top of the first rib 42 and correspondingly enhances the cooling of the first rib 42 due to the limited cooling capability of the discharged cooling air.
The preferred compound inclination of the tip notch 50 permits the first rib 42 to overhang the ramp 48, and preferably terminate with the top corner of the first rib substantially vertically or radially aligned over the outer surface of the pressure sidewall 18. In this configuration, the notch 50 has a generally L-shape with substantially normal or orthogonal legs. This configuration permits the compound notch to be readily formed in the original casting of the entire turbine blade since the mold or die material filling the notch may be readily removed from the notch in the casting process. Or, the notch may be formed by EDM, if desired.
Accordingly, the inclined first rib 42 illustrated in FIG. 3 overhangs the inclined ramp in the pressure side of the airfoil and deflects outwardly the radially outwardly migrating combustion gases for discouraging combustion gas leakage in the small gap between the tip and the surrounding shroud. The compound angled tip notch places the ramp closer to the top of the first rib and enhances film cooling thereof.
Correspondingly, the ramp apertures extend through the additional tip material provided by the ramp being inclined upwardly as well as being higher in elevation than the tip floor for providing more internal surface area within the ramp apertures in which more convection cooling may occur. The film air discharged from the ramp apertures travels a shorter distance to the overhanging first rib for obtaining improved film cooling effectiveness due to the compound notch, as well as reducing the amount of air mixing with the radially migrating combustion gases.
The compound inclination of the tip notch permits a relatively large included notch angle D which improves the ability to accurately cast the compound notch in the original blade casting, as compared with an acute included angle in this region. The typical tip shelf used in the patent identified in the Background section is horizontal or parallel with the tip floor of the blade, and if used with an inclined pressure side squealer rib it would form a relatively small acute angle therewith rendering less practical the casting of this feature in the original manufacture.
Furthermore, such a horizontal tip shelf cooperating with an inclined pressure side squealer rib would correspondingly have a relatively small fillet therebetween having a corresponding stress concentration. During rotary operation of the turbine blades, centrifugal loads would be developed in the inclined tip rib, with the centrifugal stresses generated at the base thereof near the acute shelf fillet being concentrated thereby.
In contrast, the ramp illustrated in FIG. 3 is inclined to complement the inclination of the first rib 42 and create a substantially normal or orthogonal included angle D therebetween, which normal angle is substantially greater than the corresponding acute angle if the ramp were re-configured horizontally in the form of the conventional tip shelf.
The inclined pressure side squealer rib 42 has these several advantages in aerodynamic and structural performance over the conventional horizontal tip shelf, as well as being readily formed by casting in the original blade, or subsequently by EDM. However, the suction side second squealer rib 44 is disposed downstream from the first rib 42 and is not subject to radially outward migration of the combustion gases on the pressure sidewall.
Accordingly, the second rib 44 preferably bounds the suction side of the tip cavity 46 at a substantially normal or orthogonal angle with the tip floor 40, with a fillet at the juncture therebetween.
Correspondingly, the tip floor 40 preferably includes a plurality of floor apertures 58 extending radially outwardly from the internal cooling circuit 32 and through the floor either perpendicularly or at an inclination therethrough in flow communication with the tip plenum 46. The cooling air 34 is thereby additionally channeled into the tip plenum 46 for cooling the inboard surfaces thereof including those of the first and second tip ribs, with the cooling air from the plenum then being discharged therefrom downstream over the top of the second rib 44.
In this configuration of outwardly inclined first rib 42 and normal second rib 44, the tip plenum 46 diverges radially outwardly for locally recirculating the cooling air therein, as well as for being readily castable during the original manufacture of the blade.
In the preferred embodiment illustrated in FIG. 3, the inclination angle C of the inclined ramp 48 is within the range of about 10 degrees to about 35 degrees relative to the tip floor 40, and may be about 25 degrees for example.
Correspondingly, the first rib 42 has a generally constant thickness or width with its outboard and inboard surfaces being generally parallel, with the first rib and those surfaces having an acute inclination angle A of about 65 degrees. In this way, the notch angle D may be about 90 degrees.
The inclination angle of the first rib 42 and the inclination angle of the ramp along with the respective sizes thereof determines and controls the configuration of the tip notch 50 and its performance in operation.
Compared with the vertical squealer ribs and horizontal tip shelf of the patent identified in the Background section, the inclined first rib 42 and inclined ramp 48 may be varied in value of those inclinations in a tradeoff of the various affects thereof.
The primary tradeoff occurs between the inclination angle A of the first rib 42 and the inclination angle C of the ramp 48. Increasing the inclination of the rib promotes more effective sealing performance of the rib with its cooperating shroud, but correspondingly decreases the included angle D between the rib and ramp. As that included angle decreases stress concentration at the juncture or fillet between the first rib and ramp increases, as well as increases the difficulty in casting the tip notch, leading to poor casting yield.
The inclination C of the tip ramp may be increased for increasing the included angle D between the first rib and ramp, but then the size of the tip notch 50 correspondingly decreases, which decreases the available recirculation zone for the cooling air therein which reduces the cooling effectiveness thereof along the first rib.
However, with these various interrelated geometrical features of the inclined first rib and inclined ramp, optimum values thereof may be determined for each design application depending upon the configuration of the specific turbine blade and airfoil and the intended operational environment. In the preferred embodiment disclosed above, the tip ramp 48 is preferably substantially normal to the inclined first rib 42 and may vary from perpendicular therewith within a range of about plus or minus 10 degrees, for example.
The width of the tip ramp 48 between the pressure sidewall and the first rib, including the bullnose 52 and fillet 54, may range from about 20 mils to 30 mils. The thickness of the pressure and suction sidewalls of the airfoil may be about 20 to 40 mils, with the first and second tip ribs also having this nominal dimension, which is also shared in thickness by the tip floor. The ramp apertures 56 may be about 15 mils in diameter and are preferably inclined along the length or chord direction of the tip ramp as illustrated in FIGS. 1 and 3.
The tip ramp is provided locally along the pressure sidewall immediately aft of the leading edge and terminating forward of the trailing edge where space permits and where its various benefits may be used to advantage. As the tip ramp blends at the leading edge and trailing edge along the pressure sidewall of the airfoil, the straight portion of the ramp between the bullnose 52 and fillet 54 as illustrated in FIG. 3 decreases until the arcuate curvatures of the bullnose and fillet merely join each other as the tip notch 50 and bullnose 52 disappear near the leading and trailing edges.
While there have been described herein what are considered to be preferred and exemplary embodiments of the present invention, other modifications of the invention shall be apparent to those skilled in the art from the teachings herein, and it is, therefore, desired to be secured in the appended claims all such modifications as fall within the true spirit and scope of the invention.
|Cited Patent||Filing date||Publication date||Applicant||Title|
|US4142824||Sep 2, 1977||Mar 6, 1979||General Electric Company||Tip cooling for turbine blades|
|US4893987||Dec 8, 1987||Jan 16, 1990||General Electric Company||Diffusion-cooled blade tip cap|
|US5261789||Aug 25, 1992||Nov 16, 1993||General Electric Company||Tip cooled blade|
|US5476364||Oct 27, 1992||Dec 19, 1995||United Technologies Corporation||Tip seal and anti-contamination for turbine blades|
|US5503527||Dec 19, 1994||Apr 2, 1996||General Electric Company||Turbine blade having tip slot|
|US5564902||Apr 21, 1995||Oct 15, 1996||Mitsubishi Jukogyo Kabushiki Kaisha||Gas turbine rotor blade tip cooling device|
|US5660523||Feb 3, 1992||Aug 26, 1997||General Electric Company||Turbine blade squealer tip peripheral end wall with cooling passage arrangement|
|US6039531||Mar 3, 1998||Mar 21, 2000||Mitsubishi Heavy Industries, Ltd.||Gas turbine blade|
|US6059530 *||Dec 21, 1998||May 9, 2000||General Electric Company||Twin rib turbine blade|
|US6086328||Dec 21, 1998||Jul 11, 2000||General Electric Company||Tapered tip turbine blade|
|US6164914 *||Aug 23, 1999||Dec 26, 2000||General Electric Company||Cool tip blade|
|US6224336 *||Jun 9, 1999||May 1, 2001||General Electric Company||Triple tip-rib airfoil|
|US6527514 *||Jun 11, 2001||Mar 4, 2003||Alstom (Switzerland) Ltd||Turbine blade with rub tolerant cooling construction|
|US6554575 *||Sep 27, 2001||Apr 29, 2003||General Electric Company||Ramped tip shelf blade|
|US6595749 *||Aug 28, 2001||Jul 22, 2003||General Electric Company||Turbine airfoil and method for manufacture and repair thereof|
|US6672829 *||Jul 16, 2002||Jan 6, 2004||General Electric Company||Turbine blade having angled squealer tip|
|Citing Patent||Filing date||Publication date||Applicant||Title|
|US7037075 *||Dec 1, 2003||May 2, 2006||Rolls-Royce Plc||Blade cooling|
|US7118342||Sep 9, 2004||Oct 10, 2006||General Electric Company||Fluted tip turbine blade|
|US7192250 *||Aug 3, 2004||Mar 20, 2007||Snecma Moteurs||Hollow rotor blade for the future of a gas turbine engine|
|US7270514 *||Oct 21, 2004||Sep 18, 2007||General Electric Company||Turbine blade tip squealer and rebuild method|
|US7281894 *||Sep 9, 2005||Oct 16, 2007||General Electric Company||Turbine airfoil curved squealer tip with tip shelf|
|US7287959||Dec 5, 2005||Oct 30, 2007||General Electric Company||Blunt tip turbine blade|
|US7290986||Sep 9, 2005||Nov 6, 2007||General Electric Company||Turbine airfoil with curved squealer tip|
|US7320575||Sep 28, 2004||Jan 22, 2008||General Electric Company||Methods and apparatus for aerodynamically self-enhancing rotor blades|
|US7351035 *||May 9, 2006||Apr 1, 2008||Snecma||Hollow rotor blade for the turbine of a gas turbine engine, the blade being fitted with a “bathtub”|
|US7473073 *||Jun 14, 2006||Jan 6, 2009||Florida Turbine Technologies, Inc.||Turbine blade with cooled tip rail|
|US7494319 *||Aug 25, 2006||Feb 24, 2009||Florida Turbine Technologies, Inc.||Turbine blade tip configuration|
|US7510376||Aug 25, 2005||Mar 31, 2009||General Electric Company||Skewed tip hole turbine blade|
|US7584538||Jun 21, 2007||Sep 8, 2009||General Electric Company||Method of forming a turbine blade with cooling channels|
|US7591070||Jun 21, 2007||Sep 22, 2009||General Electric Company||Turbine blade tip squealer and rebuild method|
|US7607893||Aug 21, 2006||Oct 27, 2009||General Electric Company||Counter tip baffle airfoil|
|US7645145 *||Aug 28, 2007||Jan 12, 2010||Toyota Jidosha Kabushiki Kaisha||Composite plug and electric circuit system|
|US7686578||Aug 21, 2006||Mar 30, 2010||General Electric Company||Conformal tip baffle airfoil|
|US7704047 *||Nov 21, 2006||Apr 27, 2010||Siemens Energy, Inc.||Cooling of turbine blade suction tip rail|
|US7740445||Jun 21, 2007||Jun 22, 2010||Florida Turbine Technologies, Inc.||Turbine blade with near wall cooling|
|US7857587 *||Nov 30, 2006||Dec 28, 2010||General Electric Company||Turbine blades and turbine blade cooling systems and methods|
|US7927072||Jan 22, 2007||Apr 19, 2011||Snecma||Hollow rotor blade for the turbine of a gas turbine engine|
|US8092178||Nov 28, 2008||Jan 10, 2012||Pratt & Whitney Canada Corp.||Turbine blade for a gas turbine engine|
|US8092179||Mar 12, 2009||Jan 10, 2012||United Technologies Corporation||Blade tip cooling groove|
|US8157505||May 12, 2009||Apr 17, 2012||Siemens Energy, Inc.||Turbine blade with single tip rail with a mid-positioned deflector portion|
|US8167572 *||Jul 14, 2008||May 1, 2012||Pratt & Whitney Canada Corp.||Dynamically tuned turbine blade growth pocket|
|US8172507||May 12, 2009||May 8, 2012||Siemens Energy, Inc.||Gas turbine blade with double impingement cooled single suction side tip rail|
|US8182221 *||Jul 29, 2009||May 22, 2012||Florida Turbine Technologies, Inc.||Turbine blade with tip sealing and cooling|
|US8186965||May 27, 2009||May 29, 2012||General Electric Company||Recovery tip turbine blade|
|US8246307 *||Jul 1, 2009||Aug 21, 2012||Rolls-Royce Plc||Blade for a rotor|
|US8277171||Apr 2, 2009||Oct 2, 2012||Rolls-Royce Plc||Aerofoil|
|US8313287||Jun 17, 2009||Nov 20, 2012||Siemens Energy, Inc.||Turbine blade squealer tip rail with fence members|
|US8322986 *||Jul 29, 2008||Dec 4, 2012||General Electric Company||Rotor blade and method of fabricating the same|
|US8425183||Nov 20, 2006||Apr 23, 2013||General Electric Company||Triforial tip cavity airfoil|
|US8499449||Apr 3, 2012||Aug 6, 2013||Pratt & Whitney Canada Corp.||Method for manufacturing a turbine blade|
|US8500396||Aug 21, 2006||Aug 6, 2013||General Electric Company||Cascade tip baffle airfoil|
|US8512003||Aug 21, 2006||Aug 20, 2013||General Electric Company||Tip ramp turbine blade|
|US8632311||Aug 21, 2006||Jan 21, 2014||General Electric Company||Flared tip turbine blade|
|US8727725 *||Jan 22, 2009||May 20, 2014||Florida Turbine Technologies, Inc.||Turbine vane with leading edge fillet region cooling|
|US8777567||Sep 22, 2010||Jul 15, 2014||Honeywell International Inc.||Turbine blades, turbine assemblies, and methods of manufacturing turbine blades|
|US8777572 *||Jan 9, 2012||Jul 15, 2014||Rolls-Royce Plc||Rotor blade|
|US9091177 *||Mar 14, 2012||Jul 28, 2015||United Technologies Corporation||Shark-bite tip shelf cooling configuration|
|US9228442||Apr 5, 2012||Jan 5, 2016||United Technologies Corporation||Turbine airfoil tip shelf and squealer pocket cooling|
|US9284845||Apr 10, 2014||Mar 15, 2016||United Technologies Corporation||Turbine airfoil tip shelf and squealer pocket cooling|
|US9347320||Oct 23, 2013||May 24, 2016||General Electric Company||Turbine bucket profile yielding improved throat|
|US9353632||Oct 7, 2011||May 31, 2016||Rolls-Royce Plc||Aerofoil structure|
|US9376927||Oct 23, 2013||Jun 28, 2016||General Electric Company||Turbine nozzle having non-axisymmetric endwall contour (EWC)|
|US9464528||Jun 14, 2013||Oct 11, 2016||Solar Turbines Incorporated||Cooled turbine blade with double compound angled holes and slots|
|US9528379||Oct 23, 2013||Dec 27, 2016||General Electric Company||Turbine bucket having serpentine core|
|US9551226||Oct 23, 2013||Jan 24, 2017||General Electric Company||Turbine bucket with endwall contour and airfoil profile|
|US9638041||Oct 23, 2013||May 2, 2017||General Electric Company||Turbine bucket having non-axisymmetric base contour|
|US9670784||Oct 23, 2013||Jun 6, 2017||General Electric Company||Turbine bucket base having serpentine cooling passage with leading edge cooling|
|US20040109754 *||Dec 1, 2003||Jun 10, 2004||Townes Roderick M.||Blade cooling|
|US20050063824 *||Aug 3, 2004||Mar 24, 2005||Snecma Moteurs||Hollow rotor blade for the turbine of a gas turbine engine|
|US20060051209 *||Sep 9, 2004||Mar 9, 2006||Ching-Pang Lee||Fluted tip turbine blade|
|US20060067821 *||Sep 28, 2004||Mar 30, 2006||Wadia Aspi R||Methods and apparatus for aerodynamically self-enhancing rotor blades|
|US20060088420 *||Oct 21, 2004||Apr 27, 2006||General Electric Company||Turbine blade tip squealer and rebuild method|
|US20060257257 *||May 9, 2006||Nov 16, 2006||Snecma||Hollow rotor blade for the turbine of a gas turbine engine, the blade being fitted with a "bathtub"|
|US20070059173 *||Sep 9, 2005||Mar 15, 2007||General Electric Company||Turbine airfoil curved squealer tip with tip shelf|
|US20070059182 *||Sep 9, 2005||Mar 15, 2007||General Electric Company||Turbine airfoil with curved squealer tip|
|US20070128033 *||Dec 5, 2005||Jun 7, 2007||General Electric Company||Blunt tip turbine blade|
|US20070237637 *||Aug 25, 2005||Oct 11, 2007||General Electric Company||Skewed tip hole turbine blade|
|US20070277361 *||Jun 21, 2007||Dec 6, 2007||General Electric Company||Turbine blade tip squealer and rebuild method|
|US20080044289 *||Aug 21, 2006||Feb 21, 2008||General Electric Company||Tip ramp turbine blade|
|US20080044290 *||Aug 21, 2006||Feb 21, 2008||General Electric Company||Conformal tip baffle airfoil|
|US20080044291 *||Aug 21, 2006||Feb 21, 2008||General Electric Company||Counter tip baffle airfoil|
|US20080060197 *||Jun 21, 2007||Mar 13, 2008||General Electric Company||Turbine blade tip squealer and rebuild method|
|US20080118367 *||Nov 21, 2006||May 22, 2008||Siemens Power Generation, Inc.||Cooling of turbine blade suction tip rail|
|US20080131278 *||Nov 30, 2006||Jun 5, 2008||Victor Hugo Silva Correia||Turbine blades and turbine blade cooling systems and methods|
|US20090075520 *||Aug 28, 2007||Mar 19, 2009||Toyota Jidoshia Kabushiki Kaisha||Composite Plug and Electric Circuit System|
|US20090324422 *||Aug 21, 2006||Dec 31, 2009||General Electric Company||Cascade tip baffle airfoil|
|US20100008785 *||Jul 14, 2008||Jan 14, 2010||Marc Tardif||Dynamically tuned turbine blade growth pocket|
|US20100024216 *||Jul 29, 2008||Feb 4, 2010||Donald Brett Desander||Rotor blade and method of fabricating the same|
|US20100098554 *||Jul 1, 2009||Apr 22, 2010||Rolls-Royce Plc||Blade for a rotor|
|US20100135813 *||Nov 28, 2008||Jun 3, 2010||Remo Marini||Turbine blade for a gas turbine engine|
|US20100135822 *||Nov 28, 2008||Jun 3, 2010||Remo Marini||Turbine blade for a gas turbine engine|
|US20100221122 *||Aug 21, 2006||Sep 2, 2010||General Electric Company||Flared tip turbine blade|
|US20100232979 *||Mar 12, 2009||Sep 16, 2010||Paauwe Corneil S||Blade tip cooling groove|
|US20100290919 *||May 12, 2009||Nov 18, 2010||George Liang||Gas Turbine Blade with Double Impingement Cooled Single Suction Side Tip Rail|
|US20100290920 *||May 12, 2009||Nov 18, 2010||George Liang||Turbine Blade with Single Tip Rail with a Mid-Positioned Deflector Portion|
|US20100303625 *||May 27, 2009||Dec 2, 2010||Craig Miller Kuhne||Recovery tip turbine blade|
|US20120189458 *||Jan 9, 2012||Jul 26, 2012||Rolls-Royce Plc||Rotor blade|
|US20130243596 *||Mar 14, 2012||Sep 19, 2013||United Technologies Corporation||Shark-bite tip shelf cooling configuration|
|US20140186190 *||Dec 30, 2013||Jul 3, 2014||United Technologies Corporation||Tip leakage flow directionality control|
|US20150078916 *||Sep 18, 2013||Mar 19, 2015||Honeywell International Inc.||Turbine blades with tip portions having converging cooling holes|
|US20150330229 *||Dec 30, 2013||Nov 19, 2015||United Technologies Corporation||Tip leakage flow directionality control|
|US20150330230 *||Dec 30, 2013||Nov 19, 2015||United Technologies Corporation||Tip leakage flow directionality control|
|EP2141327A2||Mar 26, 2009||Jan 6, 2010||Rolls-Royce plc||Rotor blade for a gas turbine engine|
|EP2141327A3 *||Mar 26, 2009||Jan 4, 2012||Rolls-Royce plc||Rotor blade for a gas turbine engine|
|EP2444592A1 *||Oct 6, 2011||Apr 25, 2012||Rolls-Royce plc||Rotor blade, corresponding rotor assembly and gas turbine engine|
|WO2013158194A1 *||Jan 30, 2013||Oct 24, 2013||United Technologies Corporation||Turbine airfoil tip shelf and squealer pocket cooling|
|International Classification||F01D5/18, F01D5/20|
|Cooperative Classification||F05D2250/13, Y02T50/673, F01D5/20, F01D5/187, Y02T50/676, F05D2260/202|
|European Classification||F01D5/18G, F01D5/20|
|Dec 30, 2002||AS||Assignment|
Owner name: GENERAL ELECTRIC COMPANY, NEW YORK
Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:LEE, CHING-PANG;CHERRY, DAVID GLENN;PRAKASH, CHANDER;ANDOTHERS;REEL/FRAME:013646/0326
Effective date: 20021227
|Jul 11, 2006||CC||Certificate of correction|
|Mar 14, 2008||FPAY||Fee payment|
Year of fee payment: 4
|Mar 24, 2008||REMI||Maintenance fee reminder mailed|
|Sep 23, 2011||FPAY||Fee payment|
Year of fee payment: 8
|Mar 14, 2016||FPAY||Fee payment|
Year of fee payment: 12